The standard current concept for satellite payload configuration is to layout the payload equipment on aluminium sandwich panels. A number of such panels make up an assembly called the Communications Module (CM). Thermal dissipation is typically handled by a combination of locating the dissipative units on (a) north and south facing radiator panels, and (b) internal panels (typically floors) thermally linked to the north and south facing radiators by heat pipes or pumped loops.
The disadvantages of the known current standard are:                the essentially 2-D layout delivered by a flat panel-based configuration, which leads to an inherently long rf harness length and reduced rf efficiency;        costly manufacture and assembly using potted inserts for equipment attachments to composite panels;        the flat panel structure which is reliant on the rest of the spacecraft structure or on an external handling frame to provide out of plane stiffness and overall structural integrity (leading to higher costs because of Service Module (SM)/CM interaction, CM/SM interface complexity and/or complex CM/SM Assembly, Integration and Test (AIT)).        
FIG. 1 shows schematically how a standard spacecraft payload 1 is typically used to manage the generation of waste heat. As shown, the payload 1 receives electrical power from its bus and converts the received power into transmission and reception of electromagnetic radiation. Waste heat is generated by the payload 1 during the energy conversion process.
FIG. 2 shows schematically how heat flows typically from thermally dissipative payload equipment 2 on a North/South-facing radiator panel 3 of an equatorially-orbiting conventional spacecraft. As shown in the Figure, by way of example, heat is conducted from two thermally dissipative payload equipments 2 to a face of the panel 3, and heat is then radiated from an opposing face of the panel 3 to space. The payload equipments are mechanically attached to the North/South-facing panels, which are the coldest, seeing minimal solar illumination throughout the year. Typically, the panels are between 15 mm and 25 mm thick. The panels are made from composite materials that have a high strength to mass ratio but have poor thermal conductivity. The inclination of the Earth's rotation, relative to its orbital plane around the sun, means that the North and South-facing panels alternately receive a maximum of about 60 W/m2 of solar illumination during the summer and winter seasons.
FIG. 3 is a plan view of a standard spacecraft panel layout 10, showing the typical function and interconnections between equipment. As shown in the Figure, by way of example only, the payload panel configuration has a radiating element 11 with diplexer 12 which in turn is coupled to a combination of power amplifiers 13, filter(s) 14 and splitter(s) 15 and power combiner 16. The power equipments primarily generate thermal dissipation. It is to be understood that the radiator area in use is not fully utilised because the non-thermally dissipating equipments occupy space and because lateral heat flow across the panel is poor. Note also that the panel dimensions are greater than 2 m square, typically, which results in long inefficient connections between the radiating element 11 and the power amplifiers 13.
FIG. 4 is a schematic representation of another standard spacecraft panel layout 20, showing how orthogonal heat pipes are typically embedded within the panel. In this known arrangement, heat flow across the panel is improved, as compared to that in the FIG. 2 arrangement, permitting full use of the radiator area to be utilised by dissipative equipments. Typically, the heat transport capacity of the embedded pipes is almost ten times better than required.